Detection and Assessment of Damage to Composite Structure

ABSTRACT

A method for monitoring structural integrity of a repaired aircraft component made of composite material. The method comprises: (a) placing a multiplicity of plies of repair composite material over a repair site on the component with a sensor disposed between two plies; (b) curing the plies of repair composite material so that the repair composite material, with the sensor embedded therein, is bonded to the repair site; (c) acquiring first sensor data from the sensor before a flight of the aircraft; (d) acquiring second sensor data from the sensor during or after the flight; (e) comparing the first sensor data to the second sensor data; (f) identifying differences between the first and second sensor data indicative of structural change; and (g) determining whether the identified differences indicate structural change in excess of a specified threshold. Steps (c) through (g) are performed by a computer system.

RELATED PATENT APPLICATION

This application claims the benefit, under Title 35, United States Code,§119(e), of U.S. Provisional Application No. 61/920,808 filed on Dec.26, 2013, the disclosure of which is incorporated by reference herein inits entirety.

BACKGROUND

This disclosure generally relates to systems and methods for detectingand assessing damage to composite structure. This process and method iscompatible with but not limited to composite structures that uselightning protection systems.

Modern aircraft are being designed and built with greater percentages ofcomposite materials. In some aircraft, more than 50% of the structuralcomponents are being manufactured with composite materials. Compositematerials are tough, lightweight materials. Dominating types ofcomposite materials, such as glass fibers, carbon fibers, aramid fibersor boron fibers, are combined with a coupling agent such as a resin, tocreate a product with improved or exceptional structural properties notpresent in the original materials. Composite materials are lighter andhave better mechanical and fatigue properties as compared to aluminum.However they are also less electrically conductive and provide lesselectromagnetic shielding. Reduced conductivity causes reduced currentdissipation which may result in damage when an electromagnetic effect,such as a lightning strike, occurs.

Specifically, when lightning hits an aircraft, a conductive path on theskin of the aircraft allows the electricity to travel along the skin andexit at some other location on the aircraft. Without an adequateconductive path, arcing and hot spots can occur, possibly affecting theskin. Also, the lower electrical shielding capability of compositematerials increases the lightning threat to wiring and systems withinthe aircraft,

One current mechanism used to protect composite skins on aircraftagainst lightning strike damage is to include conductive lightning skinprotection systems. Such systems may be present either in or on thecomposite skins of an aircraft. One type of system used to provide aconductive path on the aircraft is an interwoven wire fabric (IWWF).With this type of system, wires, such as phosphor-bronze wires areembedded in the top layer of the composite material nearest thewind-swept surface. Other types of systems may include the use of a thincopper foil. With an interwoven wire fabric system in the fuselage, thewires typically have a thickness range of about 0.003 to about 0.004inches. These types of wires are spaced apart from each other. A typicalspacing is around 0.010 inch in a 90-deg mesh pattern.

High-intensity electrical discharges, such as lightning strikes to acomposite material including IWWF, may result in non-compliantproperties of the IWWF within the composite material, which in turnresults in a portion of the composite material that is non-compliant.Certain portions of the non-compliant composite material may not beidentifiable by sight. The non-compliant IWWF must be replaced toprovide electromagnetic event (EME) protection for the aircraft,including removing areas with IWWF loss and replacing the removed areaswith compliant IWWF.

In addition, testing has shown that certain lightning protectivestructures tend to experience substrate microcracking and finishcracking. The microcracks tend to form due to repeated and extremetemperature, humidity, and pressure fluctuations. Microcracking occursdue to a number of factors including internal stresses from differencesin coefficient of thermal expansion, as well as from non-optimuminterface adhesion between components in composite systems.

Fiber-reinforced composite skin panels may require a localized repair toremove a portion of the panel that has been compromised. The localizedrepair includes removing the compromised portion of the panel, preparingthe area to be repaired, generally sanding surrounding composite andedge portions in a ramped or stepped manner, fabricating, bonding andcuring a composite patch that employs sufficient overlap of thecomposite material and interwoven wire fabric to ensure the transfer ofenergy from a lightning strike on the bonded repair section into thesurrounding skin panel.

Different methodologies are currently being used to inspect repairedstructures made of composite material. For example, U.S. Pat. No.7,898,246 discloses a method for non-destructive inspection of arepaired composite structure comprising interwoven wire fabric. Existingprocesses are used to validate structural repairs, but do not validateinterwoven wire fabric conductivity. In particular, they neither detectthe need for maintenance or repair of interwoven wire fabric damage norisolate or assess potential risk for EME/HIRF-related issues.

Any improvement upon the state of the art for systems and methods forinspecting and/or monitoring the health of a composite structure wouldbe beneficial, especially if such improvement could be applied to bothoriginal and repaired composite structure.

SUMMARY

The subject matter disclosed herein is directed to systems, processesand software algorithms designed to provide prognostic information andto locate and identify, assess severity, and verify maintenance andrepair of composite structure. The systems disclosed herein provideroutine information for maintenance and repair personnel and enabledamage assessments to support dispatch in the event of structural orembedded systems damage, e.g., manufacturing process impurities, repairfailure, or composite material that has been over-strained, cut, burnt,delaminated, etc.

Some processes disclosed herein provide integrated measurement and testtechniques to maintain composite structure comprised of interwoven wirefabric health throughout an airplane life cycle and to detect theoccurrence of a level of damage that could potentially lead to EME/HIRF(High-Intensity Radiated Field) related issues. In particular, thesystems disclosed herein are able to identify possible disturbances ofconductivities and/or delaminations in existing composite structure andnewly introduced composite structure due to repair to protect internalsystems from EME/HIRF-related system damage. The information acquired bythe system can be used to assess the severity of dynamic impacts, suchas those due to shockwaves or lightning strikes.

In accordance with illustrative embodiments disclosed herein, a systemand a methodology are provided for monitoring the structural integrityof a composite structure with respect to an area that has been repaired.When composite structure is repaired, there are also electricalconsiderations, due to lightning strike concerns, that are satisfied byensuring overlap of the interwoven wire fabric during patch preparation.Illustrative embodiments disclosed below employ embedded sensors,radio-frequency identification (RFID), and data extraction. Inparticular, some of the systems disclosed herein enable life cyclemonitoring of included lightning strike mitigation devices in compositerepair areas.

The composite repair systems disclosed herein comprise a sensor (sensortypes may include pressure, strain (e.g., strain gage), electricalconductivity, fiber optic, acoustical, and capacitive) which is embeddedbetween plies of a composite repair patch. Each repair patch on anaircraft can be provided with one or more sensors. In some instances, aplurality of discrete sensors are arranged at selected locations in arepair patch. In other instances, a sensor can be specifically designedto have a shape that conforms to the shape of the repair patch.

After the composite repair has been cured, the outputs of the sensor orsensors embedded in the repair site are monitored. The sensor outputsignals are then processed to identify acquired data sets indicating thepossible presence of structural damage to the repair site. In accordancewith some embodiments, the measured values output by a sensor after therepair has been completed and before the repaired parent structure(e.g., an aircraft) has been returned to service are considered abaseline.

After the repaired parent structure has been returned to service, theintegrity of the repair site can be continuously or periodicallymonitored by acquiring and processing data outputted by the sensor orsensors embedded in the repair patch. During life cycle monitoring, theoutput of each sensor is compared to a respective baseline value. Whileenvironmental conditions during service are a factor (e.g.,temperature), the sensor outputs must be processed in a manner thatremoves the effects attributable to environmental factors which wereabsent during determination of the baseline values. When the deviationof the monitored sensor output from the baseline sensor output reaches apredefined threshold (which threshold is different for every repairsituation), the patch will be repaired or replaced. The monitoringsystem has the capability to issue an alert or warning signal thatcauses the production of a visible or audible alert or warning in thecockpit, or storage of data in a memory, in response to detection of asituation wherein the monitored sensor output has deviated from thebaseline sensor output by more than a specified threshold.

One aspect of the subject matter disclosed herein a method formonitoring structural integrity of a laminated structure made ofcomposite material. The method comprises: (a) placing a sensor betweenplies of composite material which are not fully cured, the sensor beingcapable of outputting data representing a current structuralcharacteristic of surrounding composite material after the compositematerial has been cured; (b) curing the plies of composite materialwhile the sensor is in place to produce composite material having anembedded sensor; (c) after the curing step, acquiring and recordingbaseline data from the embedded sensor which represents a structuralcharacteristic of the surrounding composite material; (d) after thebaseline data has been acquired and recorded, subjecting the laminatedstructure to loads having unknown magnitudes and directions; (e)acquiring and recording post-loading data from the embedded sensor at atime subsequent to or during step (d), the post-loading datarepresenting a structural characteristic of the surrounding compositematerial; (f) processing the baseline data and post-loading data in amanner that identifies differences between the respective baseline andpost-loading data indicative of structural change in the surroundingcomposite material; and (g) determining whether the identifieddifferences indicate structural change to the surrounding compositematerial in excess of a specified threshold. The steps (e) through (g)are performed by a computer system. In some embodiments, step (f)comprises creating a baseline signature based on the baseline sensordata, creating a post-loading signature based on the post-loading sensordata, and comparing the baseline and post-loading signatures.

The foregoing method may further comprise: issuing an alert signal inresponse to a determination in step (g) that the identified differencesindicate structural change to the surrounding composite material inexcess of a specified threshold; and/or processing the post-loading datato compensate for effects due to differences in local conditions at orabout the times when steps (c) and (e) were performed.

In instances where the laminated structure comprises a parent structurehaving a repair site and a repair patch bonded to the repair site, themethod may further comprise: evaluating a current repair dispatch statusof the repair based on the results of steps (e) through (g); andspecifying an updated maintenance schedule that takes into account thecurrent repair dispatch status.

In accordance with a further aspect of the foregoing method, steps (a)through (g) are performed for each of a plurality of repairs, and theoutput from respective sensors comprises respective post-loading datafor respective repairs and respective sensor identification data forrespective sensors. When the laminated structure is part of an aircraft,that laminated structure will be subjected to loads in step (d) duringflight of the aircraft. In the latter case, the method may furthercomprise communicating the post-loading data from the sensor to acomputer system onboard the aircraft, wherein steps (e) through (g) areperformed while the aircraft is airborne; and/or communicating thepost-loading data from the sensor to a computer system on the groundafter the aircraft has landed, wherein steps (e) through (g) areperformed on the ground.

Another aspect of the subject matter disclosed herein is a systemcomprising: a parent structure made of composite material and having arepair site; a repair patch made of composite material, the repair patchbeing bonded to the parent structure at the repair site; and a sensorembedded in the repair patch. The system may further comprisenon-volatile memory and an interface unit embedded in the repair patchand electrically connected to the sensor.

A further aspect is a method for monitoring structural integrity of alaminated structure made of composite material, comprising: (a) placinga sensor between layers of composite material of a repair patch, thesensor being capable of outputting data representing a currentstructural characteristic of surrounding composite material after thecomposite material has been cured; (b) curing the composite materialwhile the repair patch is in contact with a repair site of a parentstructure made of composite material to produce a repaired parentstructure having an embedded sensor; (c) after the curing step,acquiring and recording baseline data from the embedded sensor whichrepresents a structural characteristic of the surrounding compositematerial; (d) after the baseline data has been acquired and recorded,subjecting the repaired parent structure to loads having unknownmagnitudes and directions; (e) acquiring and recording post-loading datafrom the embedded sensor at a time subsequent to or during step (d), thepost-loading data representing a structural characteristic of thesurrounding composite material; (f) processing the baseline data andpost-loading data in a manner that identifies differences between therespective baseline and post-loading data indicative of structuralchange in the surrounding composite material; and (g) determiningwhether the identified differences indicate structural change to thesurrounding composite material in excess of a specified threshold. Steps(e) through (g) are performed by a computer system.

Yet another aspect is a method for monitoring structural integrity of arepaired component of an aircraft, comprising: (a) placing amultiplicity of plies of repair composite material over a repair site onthe component with a sensor disposed between two plies; (b) curing theplies of repair composite material so that the repair compositematerial, with the sensor embedded therein, is bonded to the repairsite; (c) acquiring sensor data from the sensor before and during orafter a flight of the aircraft; (d) creating a first signature based onthe sensor data acquired before the flight; (e) creating a secondsignature based on the sensor data acquired during or after the flight;(f) comparing the first and second signatures; (g) identifyingdifferences between the first and second signatures indicative ofstructural change in the repaired aircraft component; and (h)determining whether the identified differences indicate structuralchange to the repaired aircraft component in excess of a specifiedthreshold. Steps (d) through (h) are performed by a computer system.

Other aspects of systems that monitor the structural integrity ofcomposite parts using embedded sensors are disclosed in detail andclaimed below.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagram illustrating a cross section of an aircraft skinincluding interwoven wire fabric.

FIGS. 2A-2C are diagrams showing respective types of damage to compositematerial due to lightning strikes on an aircraft. These diagramsrepresent cross-sectional views of composite material comprising alamination of plies of fiber-reinforced fabric in which the fiberorientations differ from ply to ply. The types of damage depicted are asfollows: FIG. 2A—damage due to low-energy impact; FIG. 2B—damage due tomedium-energy impact; and FIG. 2C—damage due to high-energy impact.

FIG. 3 is a diagram representing an exploded view of a damaged site on acomposite parent structure which was repaired using a patch consistingof a multiplicity of (in this simplified example, four) plies ofcomposite material, the outermost ply being IWWF.

FIG. 4 is a block diagram representing some components of a system formonitoring the structural health of a composite repair in accordancewith a wireless embodiment. The sensor can be embedded in a repair patchapplied to a repair site.

FIG. 5 is a diagram showing a plurality of idealized sensorssuperimposed on a repair site of a composite parent structure inaccordance with one embodiment. The repair patch, in which the sensorswill be embedded, is not shown in FIG. 5.

FIG. 6 is a block diagram representing some components of a sensor chipsuitable for use in the system depicted in FIG. 5.

FIG. 7 is a flowchart identifying steps of a process for acquiringinformation concerning the initial structural integrity of a repair siteon a composite structure.

FIG. 8 is a flowchart identifying steps of a process for monitoring thestructural integrity of a repair site on a composite structure.

FIG. 9 is a diagram representing an exploded view of components of asystem for measuring the electric current flowing through a pressuresensor sandwiched between plies of composite material.

FIG. 10 is a diagram showing a loop-shaped pressure sensor superimposedon a repair site of a composite parent structure in accordance withanother embodiment. The repair patch, in which the sensor will beembedded, is not shown in FIG. 10.

FIG. 11 is a diagram showing the repair site of FIG. 10 after a repairpatch has been applied. The dashed ellipse in FIG. 11 indicates theposition of the loop-shaped pressure sensor which is embedded in therepair patch and therefore hidden from view.

FIG. 12 is a flow diagram of an aircraft production and servicemethodology.

FIG. 13 is a block diagram showing systems of an aircraft.

Reference will hereinafter be made to the drawings in which similarelements in different drawings bear the same reference numerals.

DETAILED DESCRIPTION

Various embodiments of systems and methods for monitoring the structuralhealth of repaired composite material on an aircraft will now bedescribed. However, it should be appreciated that the subject matterdisclosed herein is not limited in application to repaired compositematerial, but also can be applied to original composite material,meaning that during fabrication of a composite structure, sensors(sensor types may include pressure, strain (e.g., strain gage),electrical conductivity, fiber optic, acoustical, and capacitive) can beembedded between plies of composite material at strategic locationswhere structural health monitoring is desirable. Furthermore, thecontents disclosed herein are not limited in their application tocomposite material on aircraft. Instead the contents herein haveapplication to any structure made of composite material that is intendedto comply with structural integrity specifications.

FIG. 1 is a diagram illustrating a cross section of an aircraft skinincluding interwoven wire fabric wires. In this example, the depictedstructure includes a substrate 22, an interwoven wire fabric (IWWF)layer 24, a surfacer layer 26, and a paint layer 28. The layers abovethe substrate 22 form an interwoven wire fabric component. Depending onthe implementation, the interwoven wire fabric component may onlyinclude IWWF layer 24, surfacer layer 26, and paint layer 28. In thealternative, the interwoven wire fabric component may include otherlayers in addition to or in place of the ones illustrated in thisexample.

For the purpose of illustration, it will be assumed that the substrate22 is a composite structure that serves as the skin of a fuselage. Sucha composite structure may comprise a laminate made of fiber-reinforcedplastic. As seen in FIG. 1, interwoven wire fabric layer 24 containswires 30, which provide a conductive path for electromagnetic effects,such as a lightning strike. Surfacer layer 26 provides a coating orsurface for the application of paint layer 28. The surfacer layer 26 andpaint layer 28 form a dielectric component. The dielectric component mayinclude other materials or layers, such as a sealant, a nonconductiveprimer, or some other combination of those materials along with surfacerlayer 26 and paint layer 28. The thickness of the dielectric layer abovewires 30 is important because the thickness impacts the dissipation ofelectrical energy from lightning strikes. As the dielectric thicknessincreases, greater damage may occur when lightning strikes areencountered. This type of damage is more of an unfavorable issue than asafety issue.

The damage to composite material caused by lighting strikes can berepaired using any one of a number of known methodologies. Many of theseknown repair techniques involve clean-up of the damaged site followed bythe installation of a repair patch made of composite material. Forexample, U.S. Patent Application Publ. No. 2012/0080135 discloses an insitu repair technique comprising the following steps: (1) remove paintand primer from the defective area using fine abrasive; (2) scarf (i.e.,sand) the area around the defect to a depth sufficient to clean out thedefective material and to prepare a surface for the repair plies; (3)determine the size, shape and orientation of composite repair plies,make ply templates, and kit plies (the largest repair ply should overlapat least 0.25 inch beyond the periphery of the scarf); (4) apply anadhesive ply to the repair area; (5) compact the adhesive; (6) place astack of repair plies over the compacted adhesive layer, locating andorienting plies correctly (i.e., in accordance with designspecifications) with regard to fiber direction and location tolerance;(7) compact the plies under a pressure of 1 atm; (8) subject the repairsite to a soak temperature of 220° F. for a dwell time of 30 minuteswithout applying pressure to the stack of repair plies; (9) maintainingthe soak temperature for another 30 minutes while vacuum pressure isapplied to the stack of repair plies; (9) for a patch having 35 plies orless, heat the repair site from the soak temperature to a final curetemperature of 350° F., while maintaining the vacuum pressure on thestack of repair plies, and then hold at that temperature and pressurefor 150-180 minutes to achieve full cure; (10) allow the repair site tocool; and (11) perform surface finishing as necessary.

The above-described methodology is suitable for repair of damagedcomposite material in cases where a hole is not formed and the remainingmaterial at the damaged site can be used as a mandrel for supporting therepair patch. In cases where the composite material is completelyremoved to form a hole at the damaged site, a different repairmethodology will be used involving additional tooling placed on theopposite side of the parent structure.

When repairing composite material that incorporates an IWWF layer 24(see FIG. 1), the repair patch will also include an IWWF layer.

Preferably the repair patch is configured so that the wires of the IWWFlayer of the repair patch overlap the wires of the IWWF layer of theparent structure. Nondestructive inspection techniques can be used todetermine if bonded repairs on composite panels containing interwovenwire fabrics have a sufficient overlap between the patch material andthe parent structure. The overlap width should be sufficient to allowthe transfer of energy from a lightning strike on a bonded repairsection into the parent material. The overlap width is the width of theedge of the patch material that extends over the undamaged parentmaterial. That width (typically about an inch) is preferably relativelyconstant all around the patch area.

After the aircraft has been repaired and returned to service, it ispreferable that the structural integrity of the repair be monitored atleast periodically. Lightning strikes to airplanes may occur withoutindication to the flight crew. When an airplane is struck by lightningand the strike is evident to the pilot, the pilot must determine whetherthe flight will continue to its destination or be diverted to analternate airport for inspection and possible repair. Technicians mayfind and identify lightning-strike damage by understanding themechanisms of lightning and its attachment to airplanes. Techniciansmust be aware that lightning strikes may not be reported in the flightlog because the pilots may not have known that a lightning strikeoccurred on the airplane. Having a basic understanding of lightningstrikes will assist technicians in performing effective maintenance.Thus it is desirable to monitor or check an aircraft for damage tocomposite components caused by lightning strikes at least periodicallyand preferably continuously.

FIGS. 2A-2C are diagrams showing respective types of damage to compositematerial due to lightning strikes on an aircraft. These diagramsrepresent cross-sectional views of composite material comprising alamination of plies of fiber-reinforced polymer in which the fiberorientations (e.g., 0°, ±45°, ±90) differ from ply to ply. FIG. 2Adepicts damage due to low-energy impact in the form of a matrix crack.FIG. 2B depicts damage due to medium-energy impact in the form of localfiber/matrix crushing and delaminations. FIG. 2C depicts damage due tohigh-energy impact in the form of a through penetration small damagezone, delaminations and loose fiber ends.

FIG. 3 is a diagram representing an exploded view of a damaged site 32on a composite parent structure 34 which was repaired using a patchconsisting of three repair plies 36 a-36 c of composite material (e.g.,carbon or boron fibers embedded in epoxy resin) and an outermost ply ofIWWF 38. In a known process, the composite repair plies 36 a-36 c werefabricated by determining the size, shape and orientation of thecomposite repair plies 36 a-36 c, making ply templates, laying compositematerial on the ply templates, and then curing the composite material. Arepair patch having only three plies is shown solely for the purpose ofillustrating the concept of a repair patch comprising a multiplicity ofplies. The number of plies in a composite repair patch may be fargreater than three, for example, a typical repair patch may comprise tento seventy plies.

In accordance with the systems for monitoring the structural health ofrepaired composite parts contemplated herein, one or more sensors areembedded between plies of a repair patch. The sensors should have thefollowing characteristics: (1) sensors should be compatible with thecomposite repair materials when integrated into the repair (i.e., nodelaminations, load transfer, chemical damage or electricalincompatibility); (2) sensors should be compatible with the repairprocess (i.e., able to withstand the pressures and temperatures appliedduring final cure of the installed repair patch); and (3) sensors shouldhave a sensitivity sufficient to measure the expected parameter (e.g.,pressure, stress, strain, electrical conductivity, or “goodness ofbond”) in the range of aircraft operating environments. In addition, thesensors are designed to support local data storage and wireless and/orwired data acquisition.

Many modern aircraft are provided with a Central Maintenance ComputerFunction (CMCF). The CMCF encompasses all major avionics, electrical,and mechanical systems installed on the aircraft. The CMCF collects,stores, and displays maintenance information generated by maintenancefunctionality and installed systems (e.g., member systems-initiatedtests). The CMCF has operator interface display and input devices (e.g.,multi-purpose control display units (MCDUs)).

The prior art provides airline mechanics with an electronic maintenanceterminal display that displays real-time CMCF data screens viaMultifunction Control Display Unit (MCDU) emulation. A typicalmaintenance terminal is a laptop PC comprising a cursor control device,a keyboard, an internal hard drive, a floppy diskette drive, a CD-ROMdrive, interfaces for brightness and contrast control, and a graphicaloutput printer bus. Using such a maintenance terminal, authorizedpersonnel are able to access maintenance applications that supervise theaircraft's health status.

FIG. 4 identifies some components of a system for monitoring thestructural health of a composite repair in accordance with a wirelessembodiment. A sensor 10 is embedded in a composite repair patch. Thesensor is of a type which transduces the value of a parameter into anelectrical output signal, the measured parameter being chosen becauseits value varies as a function of the degree of post-repair damageincurred by the repair site or the goodness or state of the repair.

In the embodiment depicted in FIG. 4, the embedded sensor 10 receiveselectrical power from a power supply 6 which is installed at or near therepair site. After appropriate signal conditioning (not shown in FIG. 4)of the output of sensor 10, the conditioned sensor output signal iswirelessly transmitted to either a stand-alone hand-held device 14having an antenna 16 or an avionics system 18 having an antenna 20 by atransceiver 8 having an antenna 12. The transceiver 8, which alsoreceives electrical power from the power supply 6, may be installed ator near the repair site.

The stand-alone hand-held (i.e., portable) device 14 (e.g., a laptop ortablet) may include functionality typically included in a maintenanceterminal as well as the damage detection severity assessment (DDSA)functionality disclosed herein. Current sensor data is recorded by amaintenance technician using the hand-held device 14 when the repairedaircraft is on the ground. That current sensor data will be compared tobaseline sensor data to provide an early detection report status. Thehand-held device 14 may comprise a processor programmed to comparecurrent sensor data acquired at a time after a flight by the repairedaircraft with baseline sensor data acquired at a time before that flightof the aircraft and then determine whether the differences between therespective sets of sensor data indicate that structural change greaterthan a specified threshold has occurred in the interim. In thealternative, the hand-held device 14 could download acquired currentsensor data to a CMCF or other off-board maintenance computing systemwith DDSA functionality for processing in a similar manner. Thehand-held device 14 may further comprise a display screen for displayinga visual alert when the processor determines that the sensor data isindicative of structural change greater than a specified threshold.

The avionics system 18 may be a component of the onboard CMCF thatincludes the damage detection severity assessment functionalitydisclosed herein. In particular, the avionics system 18 may include asoftware module that monitors the structural integrity of the compositerepair over time while the aircraft is airborne and provides earlydetection of delaminations and/or repair integrity within each repairsite. More specifically, the avionics system 18 may comprise a processorprogrammed to compare current sensor data acquired at a time after aflight by the repaired aircraft with baseline sensor data acquired at atime before that flight of the aircraft and then determine whether thedifferences between the respective sets of sensor data indicate thatstructural change greater than a specified threshold has occurred in theinterim. In the alternative, the avionics system 18 could downloadacquired current sensor data to an onboard CMCF for processing in asimilar manner. The avionics system 18 may further comprise a flightdeck display screen for displaying a sensory alert or maintenancemessage when the DDSA function determines that the sensor data isindicative of structural change greater than a specified threshold. Inaddition or in the alternative, the avionics system may comprise anannunciator that issues an audible alert when the processor determinesthat the sensor data is indicative of structural change greater than aspecified threshold.

Sensors can be integrated into the repair at strategic locations basedon damage and repair type analysis and depending on sensor type, repairsize and criticality. The sensors may be wired (e.g., Ethernet, USB,CANBus) or wireless (e.g., RFID, energy harvesting, WiFi) with localnon-volatile memory (NVM) to manage measurement history and status.Discrete sensors as well as sensors in the form of loops and grids, andarrays of discrete sensors are all reasonable sensor configurations.Sensors must be made of materials that integrate into the repair withoutleaving voids or causing de-bonds, e.g., possibly a tailored andcalibrated part of the repair. Sensor types may include pressure, strain(e.g., strain gage), electrical conductivity, fiber optic, acoustical,and capacitive. Some sensors may operate in a current mode, but othersmay be voltage or even acoustic or optical (e.g., embedded optical fibermay be very compatible with the repair and very sensitive to expectedpressure/strain/“goodness of bond”). Each sensor type has its own typeof signal conditioning, power and data acquisition requirements.

The sensors must be compatible with the composite repair materials whenintegrated (i.e., embedded) into the repair (e.g., no delaminations,allow load transfer, electrically compatible, no chemical damage) andcompatible with the repair process (i.e., no pressure and temperatureissues during curing of the composite repair materials), while havingsufficient sensitivity to measure the expected pressure/stress/“goodnessof bond” in the range of airplane operating environments.

In addition, the signal conditioning, power supply and data acquisitionfunctions can be hosted at different locations. For example, circuitryfor the signal conditioning, power supply and data acquisition functionscan be attached to the repair in a separate interface module,incorporated as part of a stand-alone damage detection severityassessment system; or potentially as part of a standard aircraftinterface module (such as a Remote Data Concentrator). Low-powerstand-alone applications could depend on power and data acquisition viaa wireless RFID type of interface or employ energy harvesting withlow-frequency wireless output. Continuous monitored implementationscould be powered via energy harvesting or power from vehicleinfrastructure and could be wired or wireless.

FIG. 5 is an idealized depiction showing respective positions of aplurality of sensor chips 78 a, 78 b and 78 c, each sensor beingrepresented by a respective square having a pair of intersectinginternal diagonal lines superimposed on a repair site 32 of a compositeparent structure 34 in accordance with one embodiment. The closedcontours represent the shapes of respective terraced zones that increasein depth as the zone area decreases. Respective plies of compositematerial that conform in shape to the respective terraced zones will belaid in place and then bonded to form a repair patch. The repair patch,in which the sensor chips 78 a-78 c will be embedded, is not shown inFIG. 5. The optimal position and depth of each sensor chip can bedetermined using stress analysis techniques. Preferably some sensorchips are positioned in areas where the repair is weakest, for example,along a peripheral region in which the repair patch overlies the parentstructure. In general, the number of sensors and their locations withinthe repair will depend on the type and size of the repair and stressanalysis. The sensor chips will be embedded in the composite materialwith proper grounding and protection to avoid HIRF damage effects.

In accordance with some embodiments, each embedded sensor chip may be asemiconductor chip (e.g., 3 to 5 mm square) made from materials such assilicon and selenium packed with high-temperature-resistant material.FIG. 6 is a block diagram representing some components of a sensor chipsuitable for use in the system depicted in FIG. 5. Each sensor chipcomprises a sensor 10, non-volatile memory 2 electrically connected tosensor 10 for local data storage, and an interface unit 4 electricallyconnected to sensor 10 for wireless or wired data acquisition. Theinterface unit 4 receives the data output by the sensor 10 and transmitsit to a computer system (not shown in FIG. 6) either wirelessly or via awire (not shown). Each sensor 10 may be fabricated from a materialcompatible with the composite repair materials and process (e.g.,carbon, nanotubes, glass, polymer fiber, silicon, metal foil) orconstructed as a microelectromechanical system.

The damage detection severity assessment functionality (which is asoftware application that runs on a computer system) will now bedescribed with reference to FIGS. 7 and 8.

FIG. 7 is a flowchart identifying steps of a process for acquiringinformation concerning the initial structural integrity of a repair siteon an aircraft component made of composite material, i.e., before theaircraft is put back in service. After the repair has been completed andverified (steps 50), the embedded sensors are read (step 52). In thecase of wireless communication, the data outputted by the sensorsembedded in the completed repair is transmitted to either a stand-aloneportable device or an onboard avionics (e.g., CMCF) system, aspreviously described. In the case of wired communication, the sensoroutputs are delivered to output terminals, which in turn are connectedby wires to the portable device 14 and/or the avionics system 18. Theacquired sensor data is then analyzed (step 54). Analysis andprognostics are used to characterize the repair (step 56). As usedherein, “characterization” is the act of creating a unique signaturethat is a measure of the “repair/bond goodness” of the repair. Thissignature is the basis for evaluating changes in the repair over time.After the repair has been characterized, the current repair dispatchstatus is evaluated (step 58). The evaluation of the current repairdispatch status is performed based on the “repair/bond goodness”information and its impact on dispatch of the airplane. The currentrepair dispatch status might range from “Ready to dispatch” to “Needsmaintenance within some number of cycles” to “No Dispatch untilmaintenance is performed”. In light of the state of the repair, i.e.,the “repair/bond goodness”, a schedule of required maintenance of theparticular repair is specified (step 60). All results are stored in arepair non-volatile memory and an aircraft maintenance record (steps62). Thereafter, the operator of the aircraft complies with thespecified dispatch status and the required maintenance schedule (steps64).

The aircraft operator preferably repeatedly monitors the structuralstate of the repair for the life of the aircraft in service, complyingwith the required maintenance schedule. This involves using the damagedetection severity assessment software again. Subsequent scheduled usesof the damage detection severity assessment functionality employ theinitial characterization of the repair as well as subsequent repairmeasurement data (and other model-based and empirical repairinformation) to create an updated signature. With this information, thedamage detection severity assessment can determine whether the repair isstill structurally sound or not, whether the maintenance schedule needsto be changed and in some cases that the repair is in need of immediatemaintenance before dispatch.

FIG. 8 is a flowchart identifying steps of a process for monitoring thestructural integrity of the repair site after the aircraft incorporatingthe repaired composite structure has been returned to service, exposingthe repaired composite structure to loads. During the flight of theaircraft or after the aircraft has landed, the embedded sensors are read(step 80). The acquired sensor data is then analyzed (step 82). Analysisand prognostics are used to provide an updated characterization of therepair (step 84). Then the current repair dispatch status is evaluated(step 86). In light of the state of the repair, i.e., the “repair/bondgoodness”, an updated schedule of required maintenance of the particularrepair is specified (step 88). All results are stored in a repairnon-volatile memory and an aircraft maintenance record (steps 90).Thereafter, the operator of the aircraft complies with the specifieddispatch status and the updated maintenance schedule (steps 92).

As previously disclosed, the sensors may have many differentconfigurations. For example, each repair patch may incorporate a ply inwhich a sensor is encapsulated. In accordance with one embodiment, thepressure may be an electrical resistor/conductor whoseresistance/conductivity changes as a function of the pressure beingexerted on the sensor.

FIG. 9 represents an exploded view of components of a system formeasuring pressure with a sensor 46 sandwiched between plies 40 and 42made of composite material. For example, the electrical resistance ofthe sensor 46 can be measured using measurement equipment 48 (e.g., anohmmeter).

FIG. 10 represents a loop-shaped sensor 70 superimposed on a repair site32 of a composite parent structure 34 in accordance with anotherembodiment. The repair patch, in which the sensor will be embedded, isnot shown in FIG. 10. A pair of output terminals are connected to theloop-shaped sensor 70 for data acquisition. FIG. 11 shows the repairsite of FIG. 10 after a repair patch 36 has been applied. The dashedellipse in FIG. 11 indicates the position of the loop-shaped sensor 70which is embedded in the repair patch 36 and therefore hidden from view.After the composite repair has been cured and sanded, the loop-shapedsensor 70 embedded in the repair will be electrically coupled to a powersupply.

The sensor 70 is powered and senses a baseline “goodness or state of therepair”. Then after return to service, the repair integrity isperiodically monitored/measured and evaluated. Sensor environmentalcompensation is developed and used to normalize measurements to preventfalse evaluations, which is a part of the analysis for DDSA. Significantchanges to “goodness or state of the repair” are identified duringperiodic monitoring sessions and updated signature, dispatch status andmaintenance schedule changes are reported.

The composite repair system shown in FIGS. 10 and 11 comprises aloop-shaped sensor 70 which is embedded between plies of the compositerepair patch 36. This loop-shaped sensor 70 comprises an electricallyconductive structure having an electrical conductivity that varies as afunction of the pressure exerted on that structure. For example, theloop-shaped sensor 70 may be fabricated from a material compatible withthe composite repair materials and process (e.g., carbon, nanotubes,glass, polymer fiber, silicon, metal foil) or constructed as amicroelectromechanical system. If the composite material of the repairpatch is cracked or delaminated (see FIGS. 2A-2C), this damage willchange the pressure sensed by the embedded sensor relative to a baselinepressure sensor measurements. Sensor compensation factors to correct forenvironmental differences (e.g., temperature, altitude pressure bias,cabin pressurization) between new measurements and baseline pressuremeasurements can be documented during testing and applied during the useof DDSA. This would help eliminate measurement errors due toenvironmental change. Each repair patch on an aircraft can be providedwith a respective sensor that is specifically designed to conform to theshape of the repair patch.

The monitoring system and process disclosed above may be employed in anaircraft manufacturing and service method 100 as shown in FIG. 12 forassembling and maintaining an aircraft 202 of a type depicted in FIG.13. During pre-production, exemplary method 200 may includespecification and design 204 of the aircraft 202 and materialprocurement 206. During production, component and subassemblymanufacturing 208 and system integration 210 of the aircraft 202 takesplace. Thereafter, the aircraft 222 may go through certification anddelivery 212 in order to be placed in service 214. While in service by acustomer, the aircraft 202 is scheduled for routine maintenance andservice 216 (which may also include modification, reconfiguration,refurbishment, and so on). More specifically, routine maintenance andservice 216 includes, but is not limited to, repairing damaged compositeand monitoring the structural integrity of the repaired component inaccordance with contents herein.

Each of the processes of method 200 may be performed or carried out by asystem integrator, a third party, and/or an operator (e.g., a customer).For the purposes of this description, a system integrator may includewithout limitation any number of aircraft manufacturers and major-systemsubcontractors; a third party may include without limitation any numberof venders, subcontractors, and suppliers; and an operator may be anairline, leasing company, military entity, service organization, and soon.

As shown in FIG. 13, the aircraft 202 produced by exemplary method 200may include an airframe 228 with a plurality of systems 220 and aninterior 222. Examples of high-level systems 220 include one or more ofthe following: a propulsion system 224, an electrical system 226, ahydraulic system 228, and an environmental control system 230. Anynumber of other systems may be included.

The apparatus and processes disclosed herein for monitoring thestructural integrity of composite material may be utilized duringroutine maintenance and service 216 of an aircraft 202.

While various embodiments have been described, it will be understood bythose skilled in the art that various changes may be made andequivalents may be substituted for elements thereof without departingfrom the intended scope. In addition, many modifications may be made toadapt the contents herein to a particular situation without departingfrom the scope thereof. Therefore it is intended that the claims not belimited to the particular embodiments disclosed.

As used in the claims, the term “computer system” should be construedbroadly to encompass a system having at least one computer or processor,and which may have multiple computers or processors that communicatethrough a network or bus. As used in the preceding sentence, the terms“computer” and “processor” both refer to devices having a processingunit (e.g., a central processing unit) and some form of memory (i.e.,computer-readable medium) for storing a program which is readable by theprocessing unit.

The method claims set forth hereinafter should not be construed torequire that the steps recited therein be performed in alphabeticalorder or in the order in which they are recited. Nor should they beconstrued to exclude any portions of two or more steps being performedconcurrently or alternatingly.

1. A method for monitoring structural integrity of a laminated structuremade of composite material, said method comprising: (a) placing a sensorbetween plies of composite material which are not fully cured, thesensor being capable of outputting data representing a currentstructural characteristic of surrounding composite material after thecomposite material has been cured; (b) curing the plies of compositematerial while the sensor is in place to produce composite materialhaving an embedded sensor; (c) after the curing step, acquiring andrecording baseline data from the embedded sensor which represents astructural characteristic of the surrounding composite material; (d)after the baseline data has been acquired and recorded, subjecting thelaminated structure to loads having unknown magnitudes and directions;(e) acquiring and recording post-loading data from the embedded sensorat a time subsequent to or during step (d), said post-loading datarepresenting a structural characteristic of the surrounding compositematerial; (f) processing the baseline data and post-loading data in amanner that identifies differences between the respective baseline andpost-loading data indicative of structural change in the surroundingcomposite material; and (g) determining whether the identifieddifferences indicate structural change to the surrounding compositematerial in excess of a specified threshold, wherein steps (e) through(g) are performed by a computer system.
 2. The method as recited inclaim 1, wherein step (f) comprises creating a baseline signature basedon said baseline sensor data, creating a post-loading signature based onsaid post-loading sensor data, and comparing said baseline andpost-loading signatures.
 3. The method as recited in claim 1, furthercomprising issuing an alert signal in response to a determination instep (g) that the identified differences indicate structural change tothe surrounding composite material in excess of a specified threshold.4. The method as recited in claim 1, further comprising processing thepost-loading data to compensate for effects due to differences in localconditions at or about the times when steps (c) and (e) were performed.5. The method as recited in claim 1, wherein curing the plies comprises:applying the plies to be cured as a repair patch to a portion of aparent structure to be repaired; and curing the applied plies to bondthe plies to the parent structure.
 6. The method as recited in claim 5,further comprising: evaluating a current repair dispatch status of therepair based on the results of steps (e) through (g); and specifying anupdated maintenance schedule that takes into account the current repairdispatch status.
 7. The method as recited in claim 1, wherein steps (a)through (g) are performed for each of a plurality of repairs, and theoutput from respective sensors comprises respective post-loading datafor respective repairs and respective sensor identification data forrespective sensors.
 8. The method as recited in claim 1, wherein thelaminated structure is part of an aircraft and is subjected to loads instep (d) during flight of the aircraft, said method further comprisingcommunicating the post-loading data from the sensor to a computer systemonboard the aircraft, wherein steps (e) through (g) are performed whilethe aircraft is airborne.
 9. The method as recited in claim 8, furthercomprising communicating the post-loading data from the sensor to acomputer system on the ground after the aircraft has landed, whereinsteps (e) through (g) are performed on the ground.
 10. A systemcomprising: a parent structure made of composite material and having arepair site; a repair patch made of composite material, said repairpatch being bonded to said parent structure at said repair site; and asensor embedded in said repair patch.
 11. The system as recited in claim10, further comprising non-volatile memory embedded in said repair patchand electrically connected to said sensor.
 12. The system as recited inclaim 10, further comprising an interface unit embedded in said repairpatch and electrically connected to said sensor.
 13. The system asrecited in claim 12, wherein said interface unit comprises atransceiver.
 14. The system as recited in claim 10, further comprising apower supply supported by said parent structure and connected to providepower to said sensor.
 15. A method for monitoring structural integrityof a laminated structure made of composite material, comprising: (a)placing a sensor between layers of composite material of a repair patch,the sensor being capable of outputting data representing a currentstructural characteristic of surrounding composite material after thecomposite material has been cured; (b) curing the composite materialwhile the repair patch is in contact with a repair site of a parentstructure made of composite material to produce a repaired parentstructure having an embedded sensor; (c) after the curing step,acquiring and recording baseline data from the embedded sensor whichrepresents a structural characteristic of the surrounding compositematerial; (d) after the baseline data has been acquired and recorded,subjecting the repaired parent structure to loads having unknownmagnitudes and directions; (e) acquiring and recording post-loading datafrom the embedded sensor at a time subsequent to or during step (d),said post-loading data representing a structural characteristic of thesurrounding composite material; (f) processing the baseline data andpost-loading data in a manner that identifies differences between therespective baseline and post-loading data indicative of structuralchange in the surrounding composite material; and (g) determiningwhether the identified differences indicate structural change to thesurrounding composite material in excess of a specified threshold,wherein steps (e) through (g) are performed by a computer system. 16.The method as recited in claim 15, wherein step (f) comprises creating abaseline signature based on said baseline sensor data, creating apost-loading signature based on said post-loading sensor data, andcomparing said baseline and post-loading signatures.
 17. The method asrecited in claim 15, further comprising: evaluating a current repairdispatch status based on the results of steps (e) through (g); andspecifying an updated maintenance schedule that takes into account thecurrent repair dispatch status.
 18. The method as recited in claim 15,wherein steps (a) through (g) are performed for each of a plurality ofrepairs, and the output from respective sensors comprises respectivepost-loading data for respective repairs and respective sensoridentification data for respective sensors.
 19. The method as recited inclaim 15, wherein the laminated structure is part of an aircraft and issubjected to loads during flight of the aircraft.
 20. The method asrecited in claim 19, further comprising communicating the post-loadingdata from the sensor to a computer system onboard the aircraft, whereinsteps (e) through (g) are performed while the aircraft is airborne. 21.A method for monitoring structural integrity of a repaired component ofan aircraft, comprising: (a) placing a multiplicity of plies of repaircomposite material over a repair site on the component with a sensordisposed between two plies; (b) curing the plies of repair compositematerial so that the repair composite material, with the sensor embeddedtherein, is bonded to the repair site; (c) acquiring sensor data fromthe sensor before and during or after a flight of the aircraft; (d)creating a first signature based on the sensor data acquired before theflight; (e) creating a second signature based on the sensor dataacquired during or after the flight; (f) comparing the first and secondsignatures; (g) identifying differences between the first and secondsignatures indicative of structural change in the repaired aircraftcomponent; and (h) determining whether the identified differencesindicate structural change to the repaired aircraft component in excessof a specified threshold, wherein steps (d) through (h) are performed bya computer system.
 22. The method as recited in claim 21, furthercomprising issuing an alert signal in response to a determination instep (f) that the identified differences indicate structural change tothe repaired aircraft component in excess of a specified threshold. 23.The method as recited in claim 21, further comprising processing thesensor data acquired during or after the flight prior to steps (d)through (f) to compensate for effects due to differences in localconditions at or about the times when the sensor data is acquired insteps (c).
 24. A method comprising: applying a patch to a portion of acomposite structure, the patch including plies of composite material andat least one sensor therebetween; acquiring baseline data from thesensor which represents a structural characteristic of the patch and aportion of the composite structure proximate the patch; periodicallyacquiring data from the sensor; and analyzing the periodically acquireddata and the baseline data to determine an integrity of the patch andthe composite structure proximate thereto.